Rocket propulsion is the only means that can be used beyond the atmosphere. The size of a space vessel depends essentially on its specific impulse Isp which is given by the conventional formula:ΔV=g0.Isp ln(m1/m0)  (I)in which ΔV is the speed increment, g0 is the attraction due to gravity, m1 is the launch mass, m0 is the orbital mass, and ln is the natural logarithm.
Improvements in rocket propulsion are tending to increase this specific impulse, but there are physical limits on this parameter and progress has been very slow over recent decades.
The above formula can be applied in particular to a single stage to orbit (SSTO) spacecraft that has been put into orbit. The speed increment ΔV necessary to reach low earth orbit (LEO) is about 9,000 meters per second, including losses. By convention, the residual mass of fuel can be considered as being a portion of the payload mp. The mass in orbit m0 is the sum of the empty or “dry” mass md plus the payload mass mp. With prior art rocket propulsion, it is relatively difficult to obtain sufficient payload when using a spacecraft of the SSTO type.
The present invention proposes providing a quantitative gain in the value of the specific impulse Isp of rocket propulsion while limiting the mass of the rocket engine to reasonable values.
In addition to its possible applications to space launchers starting from the Earth, the present invention can be applied to stages for propulsion in space that come into operation in orbit or starting from other planets, and which may contribute to various missions, such as, for example, a manned Earth-to-Mars mission. The levels of thrust required are then adapted to the mission in question.
In the present state of the art, there are only two solutions which enable a satisfactory ratio to be obtained between the thrust and the mass of the propulsion system. These are chemical propulsion and nuclear thermal propulsion.
Chemical propulsion is well known and is used by all launchers presently in operation. At present the highest-performance engines use multistage combustion.
The present limits on specific impulse Isp performance of chemical propulsion are due to physical limitations, the most important of which is the choice of propellant. The best known is the combination of liquid hydrogen and liquid oxygen. Small improvements can be obtained by increasing the pressure in the combustion chamber, but at the cost of increased technological difficulties. The space shuttle main engine (SSME) for the US space shuttle presents results that are the best that have presently been obtained in terms of specific impulse Isp. That is why the SSME is used as a reference when studying and comparing the embodiments proposed in accordance with the present invention.
The technical characteristics of the SSME are as follows:                mass flow rate q=468 kg/s        chamber pressure Pc=207 bars        mixture ratio=6        nozzle outlet diameter=2.39 m        expansion ratio=77        specific impulse Isp=455 s        thrust F=2090 kN        mass of engine=3 (metric) tonnes.        
The calculated pressure Pe at the nozzle outlet is about 0.176 bars.
To make comparison easier, the comparisons given below with the embodiments of the present invention have been obtained for the same mass flow rate, for the same chamber pressure, and for the same nozzle outlet pressure.
Nuclear thermal propulsion presents a specific impulse which is greater than that which can be produced by chemical propulsion. The heat generated by a nuclear reactor is transferred directly to an expelled gas which is supplied by tanks. In general, the gas is hydrogen because it has the lowest molecular mass.
Nuclear thermal propulsion was actively developed in the United States during the 1960s in the context of the NERVA program, and more recently in the context of the Timberwind program. A test installation was implemented on the ground and, over a period of 1 hour, it delivered thrust of 30 tonnes with an impulse Isp of 800 seconds. In-depth studies were also performed in Russia and tests were made on subsystems.
Programs relating to nuclear thermal propulsion are presently going slowly. One possible explanation is that in order to perform better than chemical rockets in terms of impulse Isp, it is necessary to take high risks both in programming terms and in safety terms. Specifically:                achieving impulse Isp significantly greater than that generated by present-day stages that burn liquid oxygen and liquid hydrogen implies that nuclear thermal propulsion must have the highest possible temperatures and very high pressures at the interface between the nuclear core and the outlet gases; the required performance would push technology to its limits in a portion of the engine that is critical from the safety point of view; and        it is difficult to make the internal temperature of the nuclear core uniform; as a result there is a risk of the engine being degraded because temperature margins are small compared with the technological limits of the materials.        
Furthermore, the use of a nuclear thermal engine has been envisaged until now solely for interplanetary missions, given that for an orbital mission, non-recoverable launcher debris will fall out on Earth. At the time when that type of propulsion was being studied, recoverable launchers were a long way from becoming available.
At present, and for all existing types of rocket, thrust is obtained by a gas at high pressure expanding, which gas is heated to a high temperature by a single source, whether chemical or nuclear. There are technical limits on the heating of gas, thus giving rise to limits concerning specific impulse Isp.
It will be observed that until now, the use of diversified heat sources or the introduction of heat at different locations has not been tried.
There are patents which describe a “magneto-plasma-dynamic” (MPD) technique which consists in accelerating electrons or ions present in the outlet flow.
Such acceleration is obtained by creating a force which is the result of the combined action of a current i and a magnetic field B. That type of propulsion often operates at high frequency, or in pulsed mode with pulses of duration of millisecond order.
Such a technique is described in particular in U.S. Pat. Nos. 3,173,248 (Curtiss) and 5,170,623 (Dailey).
Most devices implementing the MPD technique require special dispositions (diverging fields in the nozzle, current injection via electrodes that are generally coaxial with the flow; or else self-induced current derived from special modulation of the current flowing in the field winding).
Rocket engines implementing the MPD technique make it possible to obtain high specific impulse Isp (of the order of several thousand seconds), but the thrust obtained is very low (of the order of a few tens of N, only). Consequently, the mean weight/thrust ratio for such devices in the present prior art is most unfavorable (about 1000).